CN113356932A - Air film cooling composite hole structure for turbine blade and turbine blade - Google Patents
Air film cooling composite hole structure for turbine blade and turbine blade Download PDFInfo
- Publication number
- CN113356932A CN113356932A CN202110769505.4A CN202110769505A CN113356932A CN 113356932 A CN113356932 A CN 113356932A CN 202110769505 A CN202110769505 A CN 202110769505A CN 113356932 A CN113356932 A CN 113356932A
- Authority
- CN
- China
- Prior art keywords
- turbine blade
- dumbbell
- hole
- film
- air film
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Granted
Links
Images
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/186—Film cooling
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Landscapes
- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention belongs to the technical field of heat transfer and cooling of hot end parts of gas turbines, and relates to an air film cooling composite hole structure for a turbine blade, wherein the turbine blade comprises an air film hole plate and a thermal barrier coating arranged on the air film hole plate, a plurality of composite air film holes are formed in the air film hole plate, and a dumbbell-shaped groove is formed in the thermal barrier coating; the composite air film hole comprises a span-wise expansion section and a straight hole section which are communicated, the straight hole section is communicated with the inner cooling channel of the turbine blade, and an outlet of the span-wise expansion section is communicated with the dumbbell-shaped groove. The guide effect of the dumbbell-shaped grooves enables cooling air flow to be transversely and longitudinally diffused in the grooves, jet momentum is weakened, and the spanwise covering capability is greatly improved; in addition, the dumbbell-shaped groove changes the kidney-shaped vortex pair rotating direction, and the vortex pair develops towards two sides of the groove under the action of mainstream compression, so that cold air jet flow is better attached to the wall surface; the air film hole contains a spanwise expansion section at the outlet, so that the jet flow momentum can be further reduced, the air film diffusion in the groove is accelerated, and the cooling effect is excellent.
Description
Technical Field
The invention belongs to the technical field of heat transfer and cooling of hot end parts of gas turbines, and relates to an air film cooling composite hole structure for a turbine blade and the turbine blade.
Background
With the rapid development of aero-engines, the front inlet temperature of a turbine is continuously increased, and the inlet temperature of the turbine of an advanced engine reaches about 2000K, which is far beyond the temperature resistance limit of turbine blade materials. In order to ensure the safe operation of the turbine, various cooling technologies are adopted to cool the blades, and the cooling mainly comprises internal cooling and external cooling. External cooling is also called film cooling, namely, part of compressed air is extracted from the last stage of the compressor and directly enters the turbine by bypassing the combustion chamber, the cooled air is injected to the side surface of the high-temperature gas channel from the interior of the blade through discrete holes, the cold air is bent downstream under the action of mainstream compression and coanda effect and adheres to the vicinity of the wall surface of the blade to form a cold air film with lower temperature, the wall surface is separated from the high-temperature gas, and the radiation heat of part of the high-temperature gas to the wall surface is taken away, so that the blade is well protected. The extraction of cold air causes the reduction of system efficiency, and the realization of achieving a better air film cooling effect under the same or less cold air flow is a hot spot of the current study on the special-shaped air film holes.
At present, the discrete holes on the turbine blade mainly adopt cylindrical hole patterns, but cylindrical hole pattern jet flow is concentrated, the area between the holes is a cooled weak area, and a cooling air film cannot form good coverage on the surface of the blade. With the widespread application of thermal barrier coatings to turbine blades, a fluted film hole design has been proposed. The groove air film hole structure is characterized in that when a thermal barrier coating is sprayed on the surface of a blade, a groove is reserved at the position of an outlet of an air film hole, cooling air flows out of the air film hole, spread diffusion is formed in the groove, then the cooling air flows out of the groove to form an air film cover, and therefore the problem that the coverage of the air film in the spread direction of a cylindrical hole is insufficient is solved. The traditional transverse groove air film hole limits the diffusion of air flow in the groove in two directions, and can form a backset in the groove, so that the flow resistance is increased, and the lifting space of air film jet flow in the spreading direction diffusion capacity is reduced.
Disclosure of Invention
In order to overcome the defects in the prior art, the invention aims to provide an air film cooling composite hole structure for a turbine blade and the turbine blade, and solves the problem of low spreading capability of air film jet flow of a transverse groove air film hole.
The invention is realized by the following technical scheme:
a gas film cooling composite hole structure for a turbine blade comprises a gas film hole plate and a thermal barrier coating arranged on the gas film hole plate, wherein a plurality of composite gas film holes are formed in the gas film hole plate, and a dumbbell-shaped groove is formed in the thermal barrier coating;
the composite air film hole comprises a span-wise expansion section and a straight hole section which are communicated, the straight hole section is communicated with the inner cooling channel of the turbine blade, and an outlet of the span-wise expansion section is communicated with the dumbbell-shaped groove.
Furthermore, the dumbbell-shaped groove is formed by connecting a plurality of groove unit cells, the cross sections of the groove unit cells are two connected trapezoids, the two trapezoids share a short side, and the connecting part of the two trapezoids forms a groove throat part.
Further, the front and rear edges of the outlet of the spanwise extending section are flush with the throat of the groove.
Further, the diameter of the spanwise extending section gradually becomes smaller from the outlet to the inlet connected with the straight hole section.
Further, the diameter of the straight hole section is DhLength of L1，L1A size of (2-10) Dh(ii) a The length of the spanwise expansion section is L2，L2Size 6Dh；
The length of the groove unit cell is the same as the span-wise distance of the composite air film hole, the length of the groove unit cell is L, and the value range of L is (3-5) Dh；
The depth of the dumbbell-shaped groove is S, and the value range of S is (0.3-1.5) Dh。
Further, DhThe value range of (A) is 0.3-1.5 mm.
Further, the flow direction inclination angle of the composite air film hole is theta, and the value range of theta is 30-60 degrees.
Further, the spanwise expansion angle of the spanwise expansion section is beta, and the value range of the beta is 5-15 degrees.
Furthermore, the included angle between the side edge of the dumbbell-shaped groove and the main flow is alpha, the value range of the alpha is 60-80 degrees, and the main flow direction is perpendicular to the central axis of the dumbbell-shaped groove.
The invention discloses a turbine blade with a dumbbell-shaped groove air film cooling composite hole structure.
Compared with the prior art, the invention has the following beneficial technical effects:
the invention discloses an air film cooling composite hole structure for a turbine blade, wherein a plurality of composite air film holes are formed in an air film hole plate, and a dumbbell-shaped groove is formed in a thermal barrier coating; the composite air film hole comprises an expansion section and a straight hole section which are communicated, the air inlet is a circular straight hole section, the outlet is firstly an expansion section, then a dumbbell-shaped groove is arranged at the outlet part of the air film hole, the single hole groove is dumbbell-shaped, and the air film holes are communicated along the expansion direction of the hole row structure groove. The guide effect of the dumbbell-shaped grooves enables cooling air flow to be transversely and longitudinally diffused in the grooves, jet momentum is weakened, and the spanwise covering capability is greatly improved; in addition, the dumbbell-shaped groove changes the kidney-shaped vortex pair rotating direction, and the vortex pair develops towards two sides of the groove under the action of mainstream compression, so that cold air jet flow is better attached to the wall surface; the air film hole contains a spanwise expansion section at the outlet, so that the jet flow momentum can be further reduced, the air film diffusion in the groove is accelerated, and the cooling effect is excellent. Compared with a transverse slot hole, the dumbbell-shaped groove can realize better diffusion of cold air in the slot, reverse jet is carried out on the front edge of the groove to a certain degree, the mixing degree of the main flow and the secondary flow in the slot can be weakened, and the cold air at the downstream of the hole is better attached to the wall surface.
Furthermore, the diameter of the spanwise expanding section gradually becomes smaller from an inlet connected with the straight hole section to an outlet, and the outlet expanding type design is favorable for further improving the transverse diffusion capacity and improving the cooling effect.
Drawings
FIG. 1 is a schematic view of a dumbbell-shaped groove film-cooled composite hole configuration of a turbine blade according to the present invention;
FIG. 2 is a top view of a dumbbell-shaped groove film-cooled composite hole configuration of a turbine blade according to the present invention;
FIG. 3 is a cross-sectional view of a dumbbell-shaped groove film-cooled composite hole configuration of a turbine blade of the present invention;
FIG. 4 is a schematic view of a dumbbell-shaped groove film cooling hole watershed cell of a turbine blade of the present invention;
FIG. 5 is a schematic view of the dumbbell-shaped groove film-cooled composite hole location of a turbine blade according to the present invention.
Wherein: 1 is a gas film pore plate; 2 is a dumbbell-shaped groove front edge; 3 is the trailing edge of the dumbbell-shaped groove; 4 is a composite air film hole; 5 is the front edge of the throat part of the groove; 6 is the rear edge of the throat part of the groove; 7 is a groove unit cell; 8 is a spanwise expansion section; 9 is a straight hole section; 10 is a dumbbell-shaped groove air film hole with a pressure surface; 11 is a suction surface dumbbell-shaped groove air film hole; and 12 is an inner cooling channel.
A is main stream gas; b is a cooling air film; c is cooling air flow; d is diffused airflow in the groove; e is the groove end outflow.
Detailed Description
The invention is described in further detail below with reference to the accompanying drawings:
as shown in fig. 1-4, the invention discloses a dumbbell-shaped groove air film cooling composite hole structure of a turbine blade, the turbine blade comprises an air film hole plate 1 and a thermal barrier coating arranged on the air film hole plate 1, a plurality of composite air film holes 4 are formed in the air film hole plate 1, and dumbbell-shaped grooves are formed in the thermal barrier coating; the composite air film hole 4 comprises a span-wise expansion section and a straight hole section 9 which are communicated, the straight hole section 9 is communicated with an inner cooling channel 12 of the turbine blade, and an outlet of the span-wise expansion section is communicated with the dumbbell-shaped groove.
Specifically, as shown in fig. 2, the dumbbell-shaped groove is formed by connecting a plurality of groove unit cells 7, the cross section of each groove unit cell 7 is in two connected trapezoids, the two trapezoids share a short side, and a groove throat is formed at the joint of the two trapezoids.
As shown in figure 1, a dumbbell-shaped groove is formed in the thermal barrier coating, one side of the groove is called a dumbbell-shaped groove front edge 2, the other side of the groove is called a dumbbell-shaped groove rear edge 3, an outlet of a composite air film hole 44 is formed in the throat of the groove, and an outlet of the composite air film hole 4 is flush with the groove throat front edge 5 and the groove throat rear edge 6.
As shown in fig. 2 and 3, the length L of the dumbbell-shaped groove unit cell 7 is the same as the spanwise interval of the air film holes; wherein the composite gas film hole 4 consists of a straight hole section 9 and a spanwise expansion section 8, the length L of the straight hole section 91Length L of span-wise expansion segment 82And a spanwise divergence angle β; the depth W of the groove is controlled by the thickness of the thermal barrier coating, and the width W of the groove is the same as the distance between the front edge and the rear edge of the outlet of the composite gas film hole 4.
As shown in fig. 1, the cooling air flow C flows out along the composite air film hole 4 with a certain inclination angle θ, wherein a part of the air film hole outflow directly sprays out of the dumbbell-shaped groove, the other part forms the diffusion air flow D in the dumbbell-shaped groove, and the side outflow of the cooling air flow C and the adjacent composite air film hole 4 is converged at the end of the dumbbell-shaped groove to obtain the groove end outflow E, and finally the cooling air film B is formed. Due to the guiding effect of the dumbbell-shaped grooves, the outflow transverse momentum of the air film holes is increased, and meanwhile, the expansion diffusion capacity of jet flow in the expansion direction is further improved under the action of the expansion section 8 in the expansion direction, so that the air film coverage width is increased, and the full air film cooling is favorably realized. In addition, the geometrical structure of the dumbbell-shaped groove changes the rotation direction of kidney-shaped vortex pairs in a downstream jet flow area relative to the straight circular hole, so that the mixing degree of cold air jet flow C and main stream gas A is weakened, the wall attaching capability of a cooling air film is improved, and the cooling effect is excellent.
The design parameters of the dumbbell-shaped groove air film cooling composite hole structure are as follows:
diameter D of straight hole section 9 of composite air film hole 4hThe value range of (1) is 0.3-1.5 mm, and the length L of the straight hole section 91A size of (2-10) DhLength L of span-wise expansion segment 82Size 6DhThe spread angle beta is 5-15 degrees, and the flow direction dip angle theta of the air film hole is 30-60 degrees.
The throat part of the dumbbell-shaped groove is flush with the front edge and the rear edge of the outlet of the composite air film hole 4, the length L of the single cell 7 of the dumbbell-shaped groove is the same as the span-wise distance of the air film hole, and the value range is (3-5) DhThe included angle alpha between the side edge of the groove and the main flow direction ranges from 60 degrees to 80 degrees, and the groove depth S ranges from (0.3-1.5) Dh。
Length L of straight hole section 9 of composite air film hole 41Adjustments can be calculated to achieve larger aspect ratios.
The dumbbell-shaped groove is communicated with the outlets of the composite air film holes 4, and the number of the expansion directions of the composite air film holes is controllable; the expansion angle beta of the expanding section 8 in the gas film hole spreading direction is obtained by resolving according to the fact that the transverse length of the gas film hole outlet does not exceed 0.8 time of the length L of the dumbbell-shaped groove unit cell 7.
Example one
In the dumbbell-shaped groove air film cooling composite hole structure of the turbine blade of the embodiment, as shown in fig. 5, a pressure surface dumbbell-shaped groove air film hole 10 is formed in the pressure surface of the turbine blade, a suction surface dumbbell-shaped groove air film hole 11 is formed in the suction surface, and air is supplied through an inner cooling channel 12; a dumbbell-shaped groove is arranged at the outlet part of the composite air film hole 4 of the air film hole plate 1. Wherein the structure of the composite gas film hole 4 consists of a spanwise expanding section 8 and a straight hole section 9, and the diameter D of the straight hole section 9hThe value is 1mm, and the length L of the straight hole section 912mm, length L of the spanwise expansion section 82The size is 6mm, the expansion angle beta in the spanwise direction is 10 degrees, and the flow direction inclination angle theta of the air film hole is 35 degrees. The included angle alpha between the side edge of the dumbbell-shaped groove and the flowing direction of the main stream gas A is 70 degrees, the throat part of the groove is flush with the front edge and the rear edge of the outlet of the composite gas film hole 4, and the length L of the groove unit cell 7 is the same as the span-wise distance of the gas film hole and is 5 mm.
In this embodiment, the dumbbell-shaped groove depth S takes a value of 0.5DhI.e. 0.5mm, the value is the diameter D of the straight hole section 9 of the composite air film hole 4hHalf of the composite gas film hole 4 has a small value, and the flow direction inclination angle theta of the composite gas film hole 4 has a small value.
Example two
A pressure surface dumbbell-shaped groove air film hole 10 is formed in the pressure surface, a suction surface dumbbell-shaped groove air film hole 11 is formed in the suction surface, and air is supplied through an inner cooling channel 12; a dumbbell-shaped groove is arranged at the outlet part of the composite air film hole 4 of the air film hole plate 1. Wherein the structure of the composite gas film hole 4 consists of a spanwise expanding section 8 and a straight hole section 9, and the diameter D of the straight hole section 9hThe value is 0.5mm, and the length L of the straight hole section 912.99mm, a length L of the spanwise expansion section 82The size is 3mm, the expansion angle beta in the unfolding direction is 15 degrees, and the flow direction inclination angle theta of the air film hole is 50 degrees. The included angle alpha between the side edge of the dumbbell-shaped groove and the flowing direction of the main stream gas A is 80 degrees, the throat part of the groove is flush with the front edge and the rear edge of the outlet of the composite gas film hole 4, and the length L of the groove unit cell 7 is equal to that of the gas film holeThe spanwise spacing is the same and takes a value of 3.5mm, the groove depth S takes a value of 1.5DhI.e. 0.75 mm.
Compared with the structure in the first embodiment, the structure in the first embodiment has a larger flow direction inclination angle, the expansion angle of the composite air film hole 4 is increased for weakening the momentum in the vertical direction of the air film jet flow, meanwhile, the length L of the dumbbell-shaped groove unit cell 7 is smaller, and the outlet groove depth S and the included angle alpha between the side edge of the groove and the main flow are increased, so that the diffusion capacity of the air film in the groove is stronger, and the transverse covering capacity of the cooling air film is improved. Referring to the embodiment, in the process of adopting other implementation modes, in order to adapt to the curvature change of the turbine blade, the flow direction inclination angle of the film hole or the geometrical parameters of the dumbbell-shaped groove can be adjusted, so that the jet flow coverage is more stable, and an ideal film cooling effect is obtained.
Claims (10)
1. The gas film cooling composite hole structure for the turbine blade is characterized in that the turbine blade comprises a gas film hole plate (1) and a thermal barrier coating arranged on the gas film hole plate (1), a plurality of composite gas film holes (4) are formed in the gas film hole plate (1), and a dumbbell-shaped groove is formed in the thermal barrier coating;
the composite air film hole (4) comprises a span-wise expansion section and a straight hole section (9) which are communicated, the straight hole section (9) is communicated with an inner cooling channel (12) of the turbine blade, and an outlet of the span-wise expansion section is communicated with the dumbbell-shaped groove.
2. The film-cooling composite hole structure for the turbine blade as claimed in claim 1, wherein the dumbbell-shaped groove is formed by connecting a plurality of groove unit cells (7), the cross section of each groove unit cell (7) is in the shape of two connected trapezoids, the two trapezoids share a short side, and the connection part of the two trapezoids forms a groove throat part.
3. The film-cooled composite hole structure for a turbine blade of claim 2, wherein the leading and trailing edges of the outlet of the spanwise extending section are flush with the groove throat.
4. The film-cooled composite hole structure for a turbine blade of claim 1, wherein the diameter of the spanwise flared section is gradually reduced from the outlet to the inlet where it connects to the straight hole section (9).
5. The film-cooled composite hole structure for turbine blades according to claim 1, characterized in that the diameter of the straight hole section (9) is DhLength of L1，L1A size of (2-10) Dh(ii) a The length of the spanwise expansion section (8) is L2，L2Size 6Dh；
The length of the groove unit cell (7) is the same as the span-wise distance of the composite air film hole (4), the length of the groove unit cell (7) is L, and the value range of L is (3-5) Dh；
The depth of the dumbbell-shaped groove is S, and the value range of S is (0.3-1.5) Dh。
6. The film-cooled composite hole structure for a turbine blade of claim 5, wherein DhThe value range of (A) is 0.3-1.5 mm.
7. The film-cooled composite hole structure for a turbine blade of claim 1, wherein the flow direction inclination angle of the composite film hole (4) is θ, and θ is in the range of 30 ° to 60 °.
8. The film-cooled composite hole structure for a turbine blade of claim 1, wherein the spanwise divergent section (8) has a spanwise divergent angle β in the range of 5 ° to 15 °.
9. The film-cooling composite hole structure for the turbine blade as claimed in claim 1, wherein the included angle between the side of the dumbbell-shaped groove and the main flow is α, the value range of α is 60 ° to 80 °, and the main flow direction is perpendicular to the central axis of the dumbbell-shaped groove.
10. A turbine blade having a dumbbell-shaped grooved film-cooled composite hole structure according to any one of claims 1 to 9, wherein the dumbbell-shaped grooved film-cooled composite holes are formed in both the pressure side and the suction side of the turbine blade.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110769505.4A CN113356932B (en) | 2021-07-07 | 2021-07-07 | Air film cooling composite hole structure for turbine blade and turbine blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202110769505.4A CN113356932B (en) | 2021-07-07 | 2021-07-07 | Air film cooling composite hole structure for turbine blade and turbine blade |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN113356932A true CN113356932A (en) | 2021-09-07 |
| CN113356932B CN113356932B (en) | 2023-04-28 |
Family
ID=77538581
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202110769505.4A Active CN113356932B (en) | 2021-07-07 | 2021-07-07 | Air film cooling composite hole structure for turbine blade and turbine blade |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN113356932B (en) |
Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114687810A (en) * | 2022-03-30 | 2022-07-01 | 沈阳航空航天大学 | Turbine blade with non-uniform pre-expansion air mold hole |
Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2012202280A (en) * | 2011-03-25 | 2012-10-22 | Mitsubishi Heavy Ind Ltd | Gas turbine cooling structure |
| US8814500B1 (en) * | 2011-06-17 | 2014-08-26 | Florida Turbine Technologies, Inc. | Turbine airfoil with shaped film cooling hole |
| CN106224010A (en) * | 2016-07-18 | 2016-12-14 | 西北工业大学 | A kind of corrugated groove air film injection structure for turbo blade |
| CN106437866A (en) * | 2016-10-31 | 2017-02-22 | 中国科学院工程热物理研究所 | Discrete gas film cooling hole structure |
| CN107965353A (en) * | 2017-11-24 | 2018-04-27 | 西安交通大学 | It is a kind of that there is the jet flow groove cooling structure for improving end wall cooling effectiveness near stator blade leading edge |
| CN207829957U (en) * | 2017-12-14 | 2018-09-07 | 中国航发商用航空发动机有限责任公司 | Blade tip groove air film hole cooling structure |
| CN111578310A (en) * | 2020-04-30 | 2020-08-25 | 南京理工大学 | Air film cooling hole structure for turboshaft engine |
-
2021
- 2021-07-07 CN CN202110769505.4A patent/CN113356932B/en active Active
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JP2012202280A (en) * | 2011-03-25 | 2012-10-22 | Mitsubishi Heavy Ind Ltd | Gas turbine cooling structure |
| US8814500B1 (en) * | 2011-06-17 | 2014-08-26 | Florida Turbine Technologies, Inc. | Turbine airfoil with shaped film cooling hole |
| CN106224010A (en) * | 2016-07-18 | 2016-12-14 | 西北工业大学 | A kind of corrugated groove air film injection structure for turbo blade |
| CN106437866A (en) * | 2016-10-31 | 2017-02-22 | 中国科学院工程热物理研究所 | Discrete gas film cooling hole structure |
| CN107965353A (en) * | 2017-11-24 | 2018-04-27 | 西安交通大学 | It is a kind of that there is the jet flow groove cooling structure for improving end wall cooling effectiveness near stator blade leading edge |
| CN207829957U (en) * | 2017-12-14 | 2018-09-07 | 中国航发商用航空发动机有限责任公司 | Blade tip groove air film hole cooling structure |
| CN111578310A (en) * | 2020-04-30 | 2020-08-25 | 南京理工大学 | Air film cooling hole structure for turboshaft engine |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114687810A (en) * | 2022-03-30 | 2022-07-01 | 沈阳航空航天大学 | Turbine blade with non-uniform pre-expansion air mold hole |
| CN114687810B (en) * | 2022-03-30 | 2023-08-18 | 沈阳航空航天大学 | Turbine blade with non-uniform pre-expansion air mold holes |
Also Published As
| Publication number | Publication date |
|---|---|
| CN113356932B (en) | 2023-04-28 |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN106795771B (en) | Inner cooling system with the insertion piece for forming nearly wall cooling duct in cooling chamber in the middle part of the wing chord of gas turbine aerofoil profile | |
| CN112049690B (en) | Slot jet flow air film cooling structure for turbine end wall | |
| US7887294B1 (en) | Turbine airfoil with continuous curved diffusion film holes | |
| EP2899370B1 (en) | Turbine blade having swirling cooling channel and cooling method thereof | |
| CN113236370B (en) | Cooling structure of high-pressure moving blade of gas turbine | |
| US20180045059A1 (en) | Internal cooling system with insert forming nearwall cooling channels in an aft cooling cavity of a gas turbine airfoil including heat dissipating ribs | |
| CN106661945A (en) | Internal Cooling System With Insert Forming Nearwall Cooling Channels In An Aft Cooling Cavity Of A Gas Turbine Airfoil | |
| CN110185500A (en) | The V-type air film hole and turbo blade of turbo blade | |
| CN112682108B (en) | Turbine blade end wall structure with D-shaped micro-group air film cooling holes, method thereof and gas turbine | |
| CN106640211A (en) | Edge-blended hole structure used for air film cooling of turbine blades | |
| CN111578310A (en) | Air film cooling hole structure for turboshaft engine | |
| CN113356932A (en) | Air film cooling composite hole structure for turbine blade and turbine blade | |
| CN111706409B (en) | Corrugated air film hole with branch hole | |
| CN112483469A (en) | Rectification extension plate anti-icing structure and aviation gas turbine engine | |
| CN114687810B (en) | Turbine blade with non-uniform pre-expansion air mold holes | |
| CN111156053A (en) | Tail edge offset split structure based on gas turbine blade and cooling method | |
| CN112943379B (en) | Turbine blade separation transverse rotation re-intersection type cooling structure | |
| CN112682106B (en) | Turbine blade end wall structure with special-shaped micro-group air film cooling holes, method and gas turbine | |
| CN112922676B (en) | Internal back basin rotary cooling channel of turbine blade | |
| CN106536859A (en) | Turbine airfoil cooling system with bifurcated mid-chord cooling chamber | |
| CN210599117U (en) | Cooling structure for improving cooling effect of turbine | |
| CN114483201A (en) | Cooling hole suitable for gas turbine blade and gas turbine blade | |
| CN114109518A (en) | Turbine blade leading edge ribbed rotational flow-air film composite cooling structure | |
| CN104895620A (en) | Arrow-shaped double-hole unit structure for air film cooling | |
| CN114876583B (en) | Cooling structure of turbine movable blade trailing edge |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| GR01 | Patent grant | ||
| GR01 | Patent grant |